Conventional aircrafts are usually designed in a T-tail configuration, in which there are three tail stabilizing surfaces at the rear of the aircraft, with two horizontally oriented stabilizers mounted on either side of a vertically oriented stabilizer, resembling the shape of the letter “T” when viewed from the front or rear. An alternative configuration is the “V-tail”, also known as a “butterfly tail”, where the three tail stabilizers (two horizontal and one vertical) are replaced with two slanted stabilizers, resembling the shape of the letter “V” when viewed from the front or rear. The movable flight control surfaces differ between these two types of aircrafts. Whereas a T-tail aircraft includes “rudders” and “elevators”, for separately controlling the yaw and pitch motions, respectively, a V-tail aircraft includes “ruddervators”, which control the yaw and pitch motions jointly.
In a T-tail aircraft, the rudders are mounted on the trailing edges on either side of the vertical stabilizer (or “fin”), and the elevators are mounted on the trailing edges of each of the two horizontal stabilizers (or “tailplanes”). In a V-tail aircraft, there are two ruddervators mounted on the trailing edge of the left and right tail stabilizers, respectively. A T-tail aircraft pitches down by tilting both elevators downwards, resulting in lower pressure above each tailplane and higher pressure below, causing the tailplanes to lift and the aircraft to nose-down. Correspondingly, when both elevators are raised, the pressure is reduced below the tailplanes and raised above them, causing the aircraft to tail-down and nose-up. A V-tail aircraft pitches down by tilting the left ruddervator downward and to the left and tilting the right ruddervator downward and to the right, producing an overall tail lifting force while the resultant left and right yaw forces cancel each other out. Correspondingly, a V-tail aircraft pitches up by raising the left ruddervator upward and to the right and tilting the right ruddervator upward and to the left, producing an overall downward force on the tail stabilizers while the resultant left and right yaw forces cancel each other out.
A T-tail aircraft yaws to the right by tilting both rudders to the right, resulting in lower pressure on the left side of the fin and higher pressure to the right, causing the tail to move left and the aircraft to nose-right. Correspondingly, when both rudders are tilted to the left, the pressure is reduced on the right side of the fin and raised on the left side, causing the tail to move right and the aircraft to nose-left. A V-tail aircraft yaws to the right by tilting the left ruddervator upward and to the right while tilting the right ruddervator downward and to the right, resulting in an overall tail-rightward force (causing the aircraft to nose-right) while the resultant up and down pitch forces cancel each other out. Correspondingly, a V-tail aircraft yaws to the left by tilting the left ruddervator downward and to the left while tilting the right ruddervator upward and to the left, resulting in an overall tail-leftward force (causing the aircraft to nose-left) while the resultant up and down pitch forces cancel each other out.
In general, a V-tail aircraft has less weight and produces less drag relative to a T-tail aircraft, but requires a more complex control system to handle the flight control surfaces and also suffers reduced directional dynamic stability.
In some aircrafts, the flight control surfaces are integrally formed together with the respective tail stabilizer surfaces, rather than being formed as a separate movable trailing edge. Such a design is also referred to as a “monoblock” configuration.
Aircrafts generally have multiple control surfaces, each of which may incline or tilt about a different rotational axis, for controlling different types of aircraft motion. Reference is now made to FIGS. 1A, 1B, and 1C. FIG. 1A is a rear view schematic illustration of a V-tail aircraft 10 ruddervator, referenced 14, in a centered position about a first rotational axis, referenced 18. FIG. 1B is a rear view schematic illustration of the ruddervator 14 of FIG. 1A rotated in a clockwise direction. FIG. 10 is a rear view schematic illustration of the ruddervator 14 of FIG. 1A rotated in a counterclockwise direction.
Reference is now made to FIGS. 2A, 2B and 2C. FIG. 2A is a top view schematic illustration of a V-tail aircraft, referenced 20, with ruddervators, referenced 22 and 24, in a centered position about a second rotational axis, referenced 26. FIG. 2B is a top view schematic illustration of the V-tail aircraft 20 of FIG. 2A with ruddervators 22, 24 rotated in a first direction. In particular, both ruddervators 22, 24 are tilted toward the rear of aircraft 20 (i.e., when viewed from the top of aircraft 20, right ruddervator 22 is tilted clockwise and left ruddervator 24 is tilted counterclockwise). FIG. 2C is a top view schematic illustration of the V-tail aircraft 10 of FIG. 2A with ruddervators 22, 24 rotated in a second direction. In particular, both ruddervators 22, 24 are tilted toward the front of aircraft 20 (i.e., when viewed from the top of aircraft 20, right ruddervator 22 is tilted counterclockwise and left ruddervator 24 is tilted clockwise).
Reference is now made to FIGS. 3A, 3B and 3C. FIG. 3A is a rear view schematic illustration of a V-tail aircraft, referenced 30, with ruddervators, referenced 32 and 34, in a centered position about a third rotational axis, referenced 36. FIG. 3B is a rear view schematic illustration of the V-tail aircraft 30 of FIG. 3A with ruddervators 32, 34 rotated in a first direction. In particular, both ruddervators 32, 34 are tilted upwards (i.e., when viewed from the rear of aircraft 30, left ruddervator 32 is tilted clockwise and right ruddervator 34 is tilted counterclockwise). FIG. 3C is a rear view schematic illustration of the V-tail aircraft 30 of FIG. 3A with ruddervators 32, 34 rotated in a second direction. In particular, both ruddervators 32, 34 are tilted downwards (i.e., when viewed from the rear of aircraft 30, left ruddervator 32 is tilted counterclockwise and right ruddervator 34 is tilted clockwise).
The “angle of attack (AOA)” of an aircraft refers to the acute angle between the chord of the airfoil (i.e., aircraft wing) and the direction of undisturbed relative airflow, which is essentially the angle between the direction of the aircraft wing and the direction it is travelling. The “angle of sideslip (AOS)” refers to the angle between the aircraft centerline and the relative wind, which can be considered the directional AOA of the aircraft. An aircraft will experience stall if the aircraft exceeds a value known as the “critical angle of attack”, resulting in a rapid decrease in lift caused by a separation of airflow from the wing surface. In a stall, the wing cannot generate adequate lift to sustain level flight. The lift coefficient generally increases as a function of AOA up until a maximum point, after which it decreases dramatically. This maximum lift coefficient point corresponds to the critical AOA. A stall may occur at any pitch attitude or any airspeed, but usually occurs when the airspeed is reduced below what is known as the “unaccelerated stall speed”.
Each fixed-wing aircraft has a specific unique critical AOA at which a stall would occur. This value is usually static and predefined prior to the flight, such that the pilot and aircraft control systems can avoid reaching the critical AOA and thus avoid entering into a stall. The actual value of the critical AOA is dependent on various parameters associated with the design of the aircraft (e.g., wing profile, planform, wing aspect ratio), but is typically in the range of 8°-20°. These parameters may be influenced by the weather conditions. In particular, the temperature and humidity in the flight environment may result in the formation of ice and other forms of frozen precipitation on the surfaces of the wings, which in turn would affect the predefined critical AOA value, usually to further limit the critical AOA. Reference is now made to FIG. 4, which is a graph, generally referenced 50, showing the effect of accumulated ice on the lift coefficient of a V-tail aircraft as a function of the angle of attack. The y-axis of graph 50 represents the lift coefficient (CL), while the x-axis of graph 50 represents the angle of attack (α) in degrees. Graph 50 depicts the lift coefficient as a function of the angle of attack for V-tails with varying degrees of accumulated ice on their surfaces. Curve 52 represents a “clean V-tail”, i.e., one with no accumulated ice, while curves 54, 56 and 58, respectively represent V-tails with accumulated ice at a thickness of increasing 5% chord-wise intervals.
Some aircrafts are equipped with mechanisms for ice removal from the wings, but these mechanisms are not always completely reliable or totally effective, and may still leave a certain amount of ice. Furthermore, the weather conditions tend to change in real-time during the actual flight, and cannot be forecasted ahead of time with 100% reliability. It is possible to completely refrain from implementing flights during weather conditions that would result in ice accumulation on the wing surfaces, or to modify the flight route to mitigate the effect of these weather conditions, although these approaches are not always feasible or practical. Safety considerations should be taken into account in defining the particular critical AOA that will be utilized during the flight. In severe weather conditions such as rain, snow and ice, the aircraft must reduce loss of aerodynamic characteristics to a tolerable level and increase its aerodynamic safety margin. Unmanned aerial vehicle (UAV) aircrafts are particularly sensitive to icy weather conditions, as such aircrafts are typically not equipped with mechanisms and resources for dealing with such a scenario.
U.S. Pat. No. 5,826,834 to Potter et al, entitled “Self adaptive limiter for automatic control of approach and landing”, is directed to a fail passive flight control system for controlling the approach and landing of an aircraft. The control system includes a pitch limiter in communication with an autopilot. The limiter computes an estimated flight path angle based on vertical speed data and horizontal speed data of the aircraft. The limiter continuously computes a nominal flight path angle from the estimated flight path angle during a tracking phase of the approach/landing, until a predetermined altitude is reached and the nominal flight path angle is latched. The limiter continuously computes a nominal vertical speed based on the nominal flight path angle and horizontal speed data, and further continuously computes a vertical speed limit from the nominal vertical speed and altitude data. The limiter computes a pitch limit value from the vertical speed limit, the vertical speed, and aircraft pitch data. The autopilot limits the aircraft pitch to the pitch limit value, thus preventing the aircraft from pitching down excessively and descending below certification terrain clearance requirements.
U.S. Pat. No. 6,253,126 to Palmer, entitled “Method and apparatus for flight parameter monitoring and control”, is directed to the monitoring of aircraft flight parameters, particularly air pressures acting on various surfaces of the aircraft. According to one aspect, the skin of the aircraft is provided with small openings or ports that are connected by an air pressure conduit to pressure sensors. The ports are sensitive to air pressure changes associated with flight at different speeds. The ports are also provided with means to deter extraneous matter (e.g., water, vapor, lubrication and deicing fluids, particulates), means to prevent icing of the port, and means to decontaminate the port (e.g., a port heater and a sump volume). The air pressures are measured, recorded and stored during a first flight condition, and subsequently during a second flight condition. The measurements are compared, and utilized for deducing aerodynamic performance data (e.g., correct angle of attack and margin to stall) and determining how to control the aircraft accordingly.
U.S. Patent Application Publication No. 2009/0062973 to Caldeira et al, entitled “Stall, buffeting, low speed and high attitude projection system”, is directed to an aircraft flight control system for providing further safety controls. The aircraft control surfaces may be actuated to deploy to a certain position by a pilot interceptor (pilot input device) command. The control system monitors a set of flight parameters (e.g., angle-of-attack, angle-of-attack rate, airspeed, airspeed rate, flap position, gear position, pitch attitude, pitch rate, height above ground, ice detection) and processes the data to determine if the aircraft is operating inside a permitted envelope. If the aircraft is close to the envelope limits, the control system may bypass the pilot interceptor command to automatically position the control surfaces. The control system may protect the aircraft from scenarios such as low speeds, high attitude, stalls and buffetings.
Abzug, “V-Tail Stalling at Combined Angles of Attack and Sideslip Information”, J. Aircraft, Vol. 36, No. 4: Engineering Notes, 1999, pp. 729-731, discloses the calculation of the V-tail panel geometric angle of attack (AOA) and sideslip angle (AOS) as a function of six variables: aircraft AOA (α), aircraft AOS (β), V-tail average downwash angle (ε), V-tail average sidewash angle (σ), V-tail dihedral angle (Γ), and V-tail incidence angle (δ) for an all-moving V-tail. The calculations are valid for large AOA and AOS values to support studies of possible panel stalling. In a sample calculation of a landing approach for a 30° dihedral V-tail, the left panel would reach a stall point at an AOA of −12°, which is obtained at a right sideslip angle of 17°. The critical AOS for panel stall was found to be reduced by 3 degrees when the assumed sidewash angle is increased from 20% to 50% of the AOS. The critical AOS for panel stall was found to be reduced by 5 degrees when the downwash factor (ε0) is increased from 4 to 8. Induction from the opposite panel was found to reduce the local panel AOA of a V-tail in sideslip below those for the same AOA (i.e., raising the panel AOA at which a stall would occur), relative to the same V-tail without sideslip. Conversely, panel crossflow on a V-tail in sideslip lowers the panel AOA at which a stall would occur, relative to the same V-tail without sideslip.